1. Field of the Invention
The present invention relates generally to an air cooled turbine blade, and more specifically to trailing edge cooling of a turbine blade.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, turbine blades used in the turbine section require internal cooling to allow for higher turbine inlet temperatures that increase the efficiency of the engine. The turbine blades include a trailing edge with cooling holes that provide cooling for this region of the airfoil. in the prior art, channel flow cooling is improved with the use of pin fins or multiple impingement holes in series with a trailing edge camber line discharge to provide improved cooling capability.
FIG. 1 shows a prior art turbine blade with a trailing edge cooling circuit for a first stage turbine blade, which is exposed to the highest gas flow temperature of the blades in the turbine. The blade in FIG. 1 includes pin fins in the trailing edge cooling channel that connects a metering hole to the exit holes spaced along the trailing edge of the airfoil. Using this type of trailing edge cooling design requires a thicker trailing edge to accommodate the trailing edge cooling channel with the pin fins extending between the pressure side wall and the suction side wall. Since the side walls in the trailing edge region require a minimum thickness to provide a rigid structure of the airfoil at the trailing edge. In some turbine stage blades, this large trailing edge thickness may induce high blockage and thus reduce the stage performance.
Because of the size and space limitations, the trailing edge region of a gas turbine airfoil becomes one of the most difficult areas in the engine to cool. For a high temperature turbine airfoil cooling application, extensive trailing edge cooling is required. FIG. 2 shows another prior art first stage turbine blade with trailing edge cooling passages that makes use of a pressure side bleed for the airfoil trailing edge cooling. This type of cooling design to minimize the airfoil trailing edge thickness has been used in the airfoil trailing edge cooling for the last thirty years. Shortfalls associated with this design are the shear mixing between the cooling air and the mainstream flow as the cooling air exits from the pressure side. The shear mixing of cooling air with the mainstream flow reduces the cooling effectiveness for the trailing edge overhang and, thus, reduces the over-temperature at the airfoil trailing edge suction side location. Frequently, this over-temperature location becomes the life limiting location for the entire turbine airfoil.